This invention relates to gas turbines engines in general, and to an improved cooling scheme for airfoils.
Efficiency is a primary concern in the design of any gas turbine engine. Historically, one of the principle techniques for increasing efficiency has been to increase the gas path temperatures within the engine. Using internally cooled components made from high temperature capacity alloys has accommodated the increased temperatures. Turbine stator vanes and blades, for example, are typically cooled using compressor air worked to a higher pressure, but still at a lower temperature than that of the core gas flow passing by the blade or the vane. It will be understood that compressor bleed air for such cooling will be unavailable to support combustion in the combustor. The higher pressure provides the energy necessary to push the air through the component. A significant percentage of the work imparted to the air bled from the compressor, however, is lost during the cooling process. The lost work does not add to the thrust of the engine and negatively effects the overall efficiency of the engine. A person of skill in the art will recognize therefore, that there is a tension between the efficiency gained from higher core gas path temperatures and the concomitant need to cool turbine components and the efficiency lost from bleeding air to perform that cooling. There is, accordingly, great value in maximizing the cooling efficiency of whatever cooling air is used. Thus, to minimize any sacrifice in engine performance due to unavailability of cooling airflow to support combustion, any scheme for cooling blades and vanes must optimize the utilization of compressor bleed cooling air.
Prior art coolable airfoils, as shown in FIG. 13, typically include a plurality of internal cavities, which are supplied with cooling air. The cooling air passes through the wall of the airfoil (or the platform) and transfers thermal energy away from the airfoil in the process. The manner in which the cooling air passes through the airfoil wall is critical to the efficiency of the process. In some instances, cooling air is passed through straight or diffused cooling apertures to convectively cool the wall and establish an external film of cooling air. A minimal pressure drop is typically required across these cooling apertures to minimize the amount of cooling air that is immediately lost to the free-stream hot core gas passing by the airfoil. The minimal pressure drop is usually produced through a plurality of cavities within the airfoil connected by a plurality of metering holes. Too small a pressure drop across the airfoil wall can result in undesirable hot core gas in-flow. In all cases, the minimal dwell time in the cooling aperture as well as the size of the cooling aperture makes this type of convective cooling relatively inefficient.
One cause of ineffective cooling can be found in poor film characteristics in those applications utilizing a cooling air film to cool a wall. However, in many cases, it is desirable to establish film cooling along a wall surface. A film of cooling air traveling along the surface of the wall increases the uniformity of the cooling and insulates the wall from the passing hot core gas. A person of skill in the art will recognize, however, that film cooling is difficult to establish and maintain in the turbulent environment of a gas turbine. In most cases, air for film cooling is bled out of cooling apertures extending through the wall. The term “bled” reflects the small difference in pressure motivating the cooling air out of the internal cavity of the airfoil. One of the problems associated with using apertures to establish a cooling air film is the film's sensitivity to pressure difference across the apertures. Too great a pressure difference across an aperture will cause the air to jet out into the passing core gas rather than aid in the formation of a film of cooling air. Too small a pressure difference will result in negligible cooling airflow through the aperture, or worse, an in-flow of hot core gas. Both cases adversely affect film cooling effectiveness. Another problem associated with using apertures to establish film cooling is that cooling air is dispensed from discrete points, rather than along a continuous line. The gaps between the apertures, and areas immediately downstream of those gaps, are exposed to less cooling air than are the apertures and the spaces immediately downstream of the apertures, and are therefore more susceptible to thermal degradation.
Turbine engine blade designers and engineers are constantly striving to develop more efficient ways of cooling airfoils to prolong life of the airfoils and reduce engine operating cost. Cooling air used to accomplish this is expensive in terms of overall fuel consumption. Thus, more effective and efficient use of available cooling air in carrying out cooling of turbine airfoils is desirable not only to prolong turbine blade life but also to improve the efficiency of the engine as well, thereby again lowering engine operating cost. Consequently, there is a continuing need in the art for a cooling design that will make more effective and efficient use of available cooling air and in particular, to increase the rotor inlet temperature or decrease the cooling flow required for the same rotor inlet temperature. There is also a need in the art to facilitate entry into the new and improved cooling designs of the cooling air required to transfer heat out of the blade.